U.S. patent number 6,461,108 [Application Number 09/818,384] was granted by the patent office on 2002-10-08 for cooled thermal barrier coating on a turbine blade tip.
This patent grant is currently assigned to General Electric Company. Invention is credited to Ramgopal Darolia, Ching-Pang Lee, Robert Edward Schafrik.
United States Patent |
6,461,108 |
Lee , et al. |
October 8, 2002 |
Cooled thermal barrier coating on a turbine blade tip
Abstract
A cooling system for cooling of the squealer tip surface region
of a high pressure turbine blade used in a gas turbine engine and a
method for making a system for cooling of the squealer tip surface
region of a high pressure turbine blade used in a gas turbine
engine. The method comprises the steps of channeling apertures in a
tip cap to a diameter of about 0.004" to about 0.020" to allow
passage of cooling fluid from a cooling fluid source; applying a
bond coat of about 0.0005" to about 0.010" in thickness to the tip
cap such that the bond coat partially fills the channels; applying
a porous TBC layer of at least about 0.003" in thickness to the
bond coat, such that the porous TBC fills the channels; applying a
dense ceramic TBC layer over the porous layer; and, passing cooling
fluid from a cooling fluid source through the channel into the
porous TBC. The density of the dense TBC layer can be varied as
needed to achieve desired cooling objectives. Because the channel
exit is filled with porous TBC material, cooling fluid flows
through the porous passageways in the porous TBC layer into the
squealer tip. Although the passageways provide a plurality of
tortuous routes, the increased density of the TBC in the dense
ceramic layer provides a resistance to flow of the cooling fluid
and effectively causes the cooling fluid to more efficiently spread
through the TBC into the squealer tip before exiting into the gas
stream at the outer surface.
Inventors: |
Lee; Ching-Pang (Cincinnati,
OH), Schafrik; Robert Edward (Cincinnati, OH), Darolia;
Ramgopal (West Chester, OH) |
Assignee: |
General Electric Company
(Schenectady, NY)
|
Family
ID: |
25225420 |
Appl.
No.: |
09/818,384 |
Filed: |
March 27, 2001 |
Current U.S.
Class: |
416/96R;
29/889.1; 416/241R |
Current CPC
Class: |
F01D
5/20 (20130101); F01D 5/183 (20130101); C23C
28/00 (20130101); F01D 5/182 (20130101); F01D
5/288 (20130101); Y02T 50/67 (20130101); Y02T
50/671 (20130101); Y02T 50/676 (20130101); Y02T
50/673 (20130101); Y10T 29/49318 (20150115); Y02T
50/60 (20130101); Y02T 50/6765 (20180501) |
Current International
Class: |
F01D
5/20 (20060101); F01D 5/28 (20060101); F01D
5/14 (20060101); F01D 5/18 (20060101); C23C
28/00 (20060101); F01D 005/18 () |
Field of
Search: |
;416/97A,97R,92,231R,241A,241B,224 ;415/173.4,115
;29/889.1,899.72,889.721 ;427/454,142 |
References Cited
[Referenced By]
U.S. Patent Documents
Primary Examiner: Look; Edward K.
Assistant Examiner: McCoy; Kimya N
Attorney, Agent or Firm: Narciso; David McNess Wallace &
Nurick Santa Maria; Carmen
Parent Case Text
CROSS-REFERENCE TO RELATED APPLICATIONS
This application references co-pending applications assigned to the
assignee of the present invention, which are identified as Ser. No.
09/707,024 entitled "Multi-layer Thermal Barrier Coating with
Integrated Cooling System", and Ser. No. 09/707,027 entitled
"Transpiration Cooling in Thermal Barrier Coating", the contents of
which are incorporated herein by reference.
Claims
What is claimed is:
1. A cooling system for cooling of the squealer tip region of a
high pressure turbine airfoil used in a gas turbine engine
comprising: a superalloy tip cap; a superalloy squealer tip
extending outward in an engine radial direction from the superalloy
tip cap into a hot gas stream of the engine; at least one channel
having a first and second end, the first end terminating in an exit
orifice located on a surface of the tip cap, the second end
connecting to a cooling circuit located within a substrate, wherein
the at least one channel has a diameter to permit an effective flow
of cooling fluid; a bond coat having a thickness of about 0.0005"
to about 0.010" applied to the tip cap surface, wherein the bond
coat partially fills the exit orifice of the at least one channel;
a layer of porous thermal barrier coating (TBC) having a first
amount of predetermined porosity applied over the bond coat such
that the porous TBC substantially covers the remainder of the exit
orifice of the at least one channel; and, a layer of dense ceramic
TBC applied to the porous layer of TBC, wherein the dense TBC layer
has an amount of predetermined porosity so that it is less porous
than the porous TBC layer.
2. The cooling system of claim 1 wherein the porous TBC partially
fills the exit orifice of the at least one channel.
3. The cooling system of claim 1 wherein the porous TBC completely
fills the exit orifice of the at least one channel.
4. The cooling system of claim 1 wherein the at least one channel
has a cross-sectional area equivalent to a diameter of about 0.004"
to about 0.020".
5. The cooling system of claim 1 wherein the bond coat has a
thickness of about 0.002".
6. The cooling system of claim 1 wherein the bond coat is an
aluminide selected from the group consisting of NiAl and PtAl and
combinations thereof.
7. The cooling system of claim 1 wherein the bond coat is a
MCrAl(X) where M is an element selected from the group consisting
of Fe, Co and Ni and X is an element selected from the group
consisting of gamma prime formers, solid solution strengtheners,
grain boundary strengtheners, reactive elements and combinations
thereof.
8. The cooling system of claim 1 wherein the layer of porous TBC
having a first amount of predetermined porosity has a thickness of
at least about 0.003" so that the porous TBC fills the remainder of
the exit orifice and a cooling fluid can pass substantially
unrestricted through the porous layer.
9. The cooling system of claim 1 wherein the layer of porous TBC
having a first amount of predetermined porosity has a thickness of
from about 0.003" or less so that the exit orifice is not
completely filled and a cooling fluid can pass substantially
unrestricted through the porous layer.
10. The cooling system of claim 1 wherein the dense ceramic layer
is selected from the group consisting of yttria-stabilized
zirconia, zirconia modified by refractory oxides, Al.sub.2 O.sub.3,
oxides formed from Group IV, V and VI elements and oxides modified
by Lanthanide Series elements.
11. The cooling system of claim 1 further including a cooling fluid
supplied from the cooling circuit, whereby the cooling fluid is
diffused and flows through the layer of porous TBC.
12. The cooling system of claim 1 further including at least one
opening extending through the porous TBC layer and opening onto an
outer surface.
13. The cooling system of claim 1 further including a TBC layer
applied to at least one of the group consisting of a pressure side
and a suction side of an airfoil.
14. The cooling system of claim 1 wherein the dense ceramic layer
is applied to a thickness of from about 0.002"-0.020".
15. The cooling system of claim 12 wherein the dense ceramic layer
is applied to a thickness of from about 0.002"-0.003".
16. A method for cooling of the squealer tip region of a high
pressure turbine blade used in a gas turbine engine comprising the
steps of: channeling apertures having a diameter of about 0.004" to
about 0.020" in a tip cap of the turbine blade to allow passage of
cooling fluid from a cooling fluid source to a surface of the tip
cap; forming a bond coat having a thickness in the range of about
0.0005" to about 0.010" to an outer surface of the tip cap and at
least adjacent squealer tip walls such that the bond coat coats
walls of the apertures formed in substrate material near exit
orifices at the tip cap surface; applying a porous thermal barrier
coating (TBC) layer having a first preselected density having a
thickness of at least about 0.003" over the formed bond coat, such
that the TBC covers the tip cap outer surface and adjacent squealer
tip walls and at least partially fills the remainder of the exit
orifices; applying a dense ceramic layer over the porous TBC layer,
wherein the dense ceramic layer has a second preselected density
that is more dense than the first preselected density of the porous
TBC layer; and, passing cooling fluid from the cooling fluid source
through the apertures in the tip cap, into and through the porous
TBC layer.
17. The method of claim 16 wherein the apertures are channeled in
the tip cap by laser drilling.
18. The method of claim 16 wherein the apertures have a
substantially circular cross-section.
19. The method of claim 16 wherein the bond coat is applied a
thickness of about 0.002".
20. The method of claim 16 further comprising the step of forming
openings extending through the porous TBC layer and opening onto an
outer surface.
21. The method of claim 16 further comprising the step of applying
a TBC layer to at least one of the group consisting of a pressure
side and a suction side of an airfoil.
22. A cooling system for cooling of the squealer tip surface region
of a high pressure turbine blade used in a gas turbine engine
formed by the method of claim 16.
Description
FIELD OF THE INVENTION
This invention relates generally to gas turbine engines, and in
particular, to a cooled flow path surface region of a turbine blade
tip.
BACKGROUND OF THE INVENTION
In gas turbine engines, for example, aircraft engines, air is drawn
into the front of the engine, compressed by a shaft-mounted
rotary-type compressor, and mixed with fuel. The mixture is burned,
and the hot exhaust gases are passed through a turbine mounted on a
shaft. The flow of gas turns the turbine, which turns the shaft and
drives the compressor and fan. The hot exhaust gases flow from the
back of the engine, driving it and the aircraft forward.
During operation of gas turbine engines, the temperatures of
combustion gases may exceed 3,000.degree. F., considerably higher
than the melting temperatures of the metal parts of the engine,
which are in contact with these gases. Operation of these engines
at gas temperatures that are above the melting temperatures of the
metal components is a well established art, and depends in part on
supplying a cooling air to the outer surfaces of the metal parts
through various methods. The metal parts of these engines that are
particularly subject to high temperatures, and thus require
particular attention with respect to cooling, are metal parts
forming combustors and parts located aft of the combustor including
turbine blades, turbine vanes and exhaust nozzles.
The hotter the turbine inlet gases, the more efficient is the
operation of the jet engine. There is thus an incentive to raise
the turbine inlet gas temperature. However, the maximum temperature
of the turbine inlet gases is normally limited by the materials
used to fabricate the turbine vanes and turbine blades of the
turbine. In current engines, the turbine vanes and blades are made
of nickel-based superalloys, and can operate at metal surface
temperatures of up to 2100.degree.-2200.degree. F.
The metal temperatures can be maintained below melting levels with
current cooling techniques by using a combination of improved
cooling designs and insulating thermal barrier coatings (TBCs). For
example, with regard to the metal blades and vanes employed in
aircraft engines, some cooling is achieved through convection by
providing passages for flow of cooling air internally within the
blades so that heat may be removed from the metal structure of the
blade by the cooling air. Such blades essentially have intricate
serpentine passageways within structural metal forming the cooling
circuits of the blade.
Small internal orifices have also been devised to direct this
circulating cooling air directly against certain inner surfaces of
the airfoil to obtain cooling of the inner surface by impingement
of the cooling air against the surface, a process known as
impingement cooling. In addition, an array of small holes extending
from the hollow core through the blade shell can provide for
bleeding cooling air through the blade shell to the outer surface
where a film of such air can protect the blade from direct contact
with the hot gases passing through the engines, a process known as
film cooling.
In another approach, a TBC is applied to the turbine blade
component, which forms an interface between the metallic component
and the hot gases of combustion. The TBC includes a ceramic coating
that is applied to the external surface of metal parts within
engines to impede the transfer of heat from hot combustion gases to
the metal parts, thus insulating the component from the hot
combustion gas. This permits the combustion gas to be hotter than
would otherwise be possible with the particular material and
fabrication process of the component. TBCs have also been used in
combination with film cooling techniques wherein an array of fine
holes extends from the hollow core through the TBC to provide
cooling air onto the outer surface of the TBC.
Certain designs of airfoil tips utilize film cooling techniques.
Film cooling is achieved by passing cooling air through discrete
film cooling holes, typically ranging from 0.015" to about 0.030"
in hole diameters. The film cooling holes are typically drilled
with laser or EDM or ES machining. Due to mechanical limitations,
each film hole has an angle ranging from 20.degree. to 90.degree.
relative to the external surface. Therefore, each film jet exits
from the hole with a velocity component perpendicular to the
surface. Because of this vertical velocity component and a complex
aerodynamic flow circulation near the tip of a turbine blade,
commonly referred to as the "squealer tip", each film jet will have
a tendency to lift or blow off from the external surface and mix
with the hot exhaust gases, resulting in poor film cooling
effectiveness in the area surrounding the squealer tip.
TBCs are well-known ceramic coatings, for example,
yttrium-stabilized zirconia (YSZ). Ceramic TBCs usually do not
adhere well directly to the superalloys used in the substrates.
Therefore, an additional metallic layer called a bond coat is
placed between the substrate and the thermal barrier coating. The
bond coat may be made of a nickel-containing overlay alloy, such as
a MCrAlX, or other compositions more resistant to environmental
damage than the substrate, or alternatively, the bond coat may be a
diffusion nickel aluminide or platinum aluminide, whose surface
oxidizes to a protective aluminum oxide scale that provides
improved adherence to the ceramic top coatings. The bond coat and
the overlying TBC are frequently referred to as a thermal barrier
coating system.
Multi layer coatings are known in the art. For example, U.S. Pat.
No. 5,846,605 to Rickerby et al. is directed to a coating having a
plurality of alternate layers having different structures that
produce a plurality of interfaces. The interfaces provide paths of
increased resistance to heat transfer to reduce thermal
conductivity. A bond coat overlying a metallic substrate is bonded
to a TBC. The TBC comprises a plurality of layers, each layer
having columnar grains, the columnar grains in each layer extending
substantially perpendicular to the interface between the bond coat
and metallic substrate. The structure is columnar to ensure that
the strain tolerance of the ceramic TBC is not impaired. The
difference in structure of the layers is the result of variations
in the microstructure and/or density/coarseness of the columnar
grains of the ceramic.
U.S. Pat. No. 5,705,231 to Nissley et al. is directed to a
segmented abradable ceramic coating system having enhanced
abradability and erosion resistance. A segmented abradable ceramic
coating is applied to a bond coat comprising three ceramic layers
that are individually applied. There is a base coat foundation
layer, a graded interlayer, and an abradable top layer. The coating
is characterized by a plurality of vertical microcracks.
U.S. Pat. No. 4,503,130 to Bosshart et al. is directed to coatings
having a low stress to strength ratio across the depth of the
coating. Graded layers of metal/ceramic material having increasing
ceramic composition are sequentially applied to the metal substrate
under conditions of varied substrate temperature. Excessive
stresses induced by differential strains between the layers is
avoided. The effect of substrate temperature control and the
differing coefficients of thermal expansion between materials of
successive layers are matched to achieve the desired result.
U.S. Pat. No. 6,045,928 to Tsantrizos et al. is directed to a TBC
comprising an MCrAlY bond coat and a dual constituent ceramic
topcoat. The topcoat comprises a monolithic zirconia layer adjacent
to the bond coat, a monolithic layer of calcia-silica representing
the outer surface of the TBC and a graded interface between the two
to achieve good adhesion between the two constituents to achieve an
increased thickness of the topcoat, thereby, providing for a
greater temperature drop across the TBC system. As used by
Tsantrizos et al., monolithic refers to a uniform composition of a
layer, while a graded interface refers to a layer having a changing
composition from one monolithic composition to the other monolithic
composition.
U.S. Pat. No. 4,576,874 to Spengler et al. is directed to a coating
to increase resistance to spalling and corrosion. The coating is
not intended to be a thermal barrier coating. A porous ceramic is
applied over a MCrAlY bond coat and a dense ceramic is then applied
over the porous ceramic. The porous portion is a transition zone to
allow for differences in thermal expansion and provides little
thermal insulation.
Improved environmental resistance to destructive oxidation and hot
corrosion is desirable. In addition, the alloying elements of the
bond coat interdiffuse with the substrate alloy, changing the
composition of the protective outer layer so that the walls of the
turbine airfoils are consumed. This loss of material reduces the
load carrying capability of the airfoil, thereby limiting blade
life. This interdiffusion can also reduce environmental resistance
of the coating. This interdiffusion and its adverse effects can be
reduced by controlling the temperature of the component in the
region. of the bond coat/substrate interface.
Thus, there is an ongoing need for an improved thermal barrier
coating system, especially surrounding the squealer tip, wherein
the environmental resistance and long-term stability of the thermal
barrier coating system is improved so that higher engine
efficiencies can be achieved. The bond coat temperature limit is
critical to the TBC's life and has had an upper limit of about
2100.degree. F. Once the bond coat exceeds this temperature, the
coating system will quickly deteriorate, due to high temperature
mechanical deformation and oxidation, as well as interdiffusion of
elements with the substrate alloy. The coating system can separate
from the substrate exposing the underlying superalloy component to
damage from the hot gasses.
In particular, the squealer tip is the most difficult location to
cool in a turbine blade. The squealer tip is located away from the
convection cooling in the center of the blade, and the complex
aerodynamic flow field near the squealer tip makes film cooling
very inefficient. This inefficient cooling results in tip
deterioration much earlier than desired, and requires tip repairs
after relatively short time in-service to recover the tip clearance
for better turbine efficiency.
As described above, to be more effective in permitting the
attainment of higher engine operating temperatures, a TBC requires
active cooling on the backside of the region being cooled. A TBC
has not been considered for use in the squealer tip region or
airfoils partly because of physical constraints and partly because
no backside cooling was available to take advantage of the
capabilities of the TBC by removing heat, thereby preventing a
build-up of temperature in this region. During the airfoil
manufacturing process, to prevent application of TBC coating in
this area, the squealer tip and cap have usually been masked during
the TBC coating process. However, recently, in order to reduce the
manufacturing costs associated with the time consuming process of
masking, the TBC coating application has been extended to cover the
squealer tip, thus avoiding the masking process. The consequence of
the presence of the TBC in this region has been that the higher
temperatures experienced in this region because of a lack of
cooling causes the TBC to spall from the backside of the squealer
tip region after several cycles of engine operation, resulting in
the same general configuration that occurs when masking is
performed to prevent the deposition of the TBC. However, since it
has become routine to apply a TBC to this region, it would
therefore be advantageous to take advantage of the presence of the
TBC by further improving squealer tip cooling by intentionally
incorporating an effective TBC coating system to extend squealer
tip life, which can be accomplished by providing adequate cooling
of the region.
What is needed are improved designs that will allow a turbine
engine blade squealer tip to run at higher operating temperatures,
thus improving engine performance without the need for additional
cooling air for the blade. It is also desirable to have a system
that can take advantage of the thermal insulation provided by TBC.
The present invention fulfills this need, and further provides
related advantages.
SUMMARY OF THE INVENTION
The present invention provides a method for cooling the squealer
tip region of a high pressure turbine blade used in a gas turbine
engine comprising the steps of channeling apertures in a substrate
to a diameter sufficient to allow passage of cooling fluid from a
cooling fluid source; applying a sufficiently thick bond coat to
the substrate such that the bond coat partially fills the
apertures; applying a porous TBC layer to the bond coat, such that
the TBC partially or completely fills the apertures; and applying a
dense ceramic layer that is more dense than the porous TBC layer on
top of the porous TBC layer. Optionally, conventional TBC can be
applied on the concave (pressure side) and convex (suction side) of
the airfoil surface.
In this manner, cooling fluid passes from a cooling fluid source
through a channel aperture adjacent to the squealer tip into the
porous TBC. Because the channel aperture is filled with porous TBC
material, cooling fluid flows through the porous passageways into
the porous TBC layer, continuing to flow between the bond coat and
the dense coat, exiting to the squealer tip or any locations inside
the tip cavity. In this manner, cooling fluid is directed to the
squealer tip, previously unobtainable using known methods.
The present invention further comprises both the cooled blade and
squealer tip region formed by the foregoing methods and the blade
and squealer tip with the ceramic layers for cooling the squealer
tip.
One advantage of the present invention is that the passageways
provide a plurality of tortuous routes, whereby the increased
density of the TBC in the dense outer layer having reduced porosity
provides a resistance to flow of the cooling fluid and effectively
causes the cooling fluid to more efficiently spread through the
porous TBC before exiting at the outer surface.
Another advantage of the present invention is that the
multi-layered TBC system forms cooling paths which allow cooling
air from the interior of the blade to flow from tip holes adjacent
to the squealer tip into the porous ceramic layer flowing between
the bond coat and the dense coat , thereby flowing to the tip or
any locations inside the tip cavity, providing efficient active
convection cooling for both the squealer tip substrate and the bond
coat by allowing heat to be removed from the squealer tip.
By removing heat from this region, the integrity of the bond coat
can be maintained at higher engine firing temperatures, resulting
in a more efficient usage of cooling fluid than that of the prior
art to achieve a higher turbine engine efficiency and performance
while improving squealer tip service life.
Still another advantage of the present invention is that the
cooling channel exit apertures, being at least partially filled
with porous TBC, have more flow resistance than open apertures and,
therefore, provide a more uniform cooling flow distribution
compared to unfilled conduits that transfer the same amount of
cooling fluid, resulting in more efficient heat transfer.
Other features and advantages of the present invention will be
apparent from the following more detailed description of the
preferred embodiment, taken in conjunction with the accompanying
figures which illustrate, by way of example, the principles of the
invention.
BRIEF DESCRIPTION OF THE DRAWINGS
FIG. 1 is a representation of a form of existing art cooling of a
turbine blade tip.
FIG. 2 is a representation of a form of the multi-layer ceramic
coating of the present invention.
DETAILED DESCRIPTION OF THE INVENTION
Substrate materials often used in turbine parts or airfoils for
aircraft engines and power generation equipment may include nickel,
cobalt, or iron based superalloys. The alloys may be cast or
wrought superalloys. Examples of such substrates are GTD-111,
GTD-222, Rene' 80, Rene' 41, Rene' 125, Rene' 77, Rene' N4, Rene'
N5, Rene' N6, 4.sup.th generation single crystal superalloy, MX-4,
Hastelloy X, and cobalt-based HS-188.
As shown in FIG. 1, a known squealer tip design utilizes cavity
purge holes through the tip cap and pressure side film cooling
holes that do not permit cooling fluid to flow to the squealer tip.
The squealer tip and tip cap do not include state-of-the-art
insulation such as thermal barrier coating systems, as the
available cooling methods are inadequate to remove heat from these
portions of the airfoil. The method of the present invention
results in, for example, the airfoil blade tip shown in FIG. 2,
which provides convection cooling and permits an effective use of
well-known insulation. Optionally, TBC coating may be applied to
either the pressure side, the suction side or both sides as well as
to the tip cap.
A plurality of small channels 6, having a cross-sectional area
equivalent to that obtained from a circular hole having a diameter
of from about 0.004" to about 0.020", and preferably from a
circular hole having a diameter of from about 0.004" to about
0.008", the cross-section having a preselected configuration,
preferably a substantially circular cross-section, are drilled,
such as by laser beam, electrical discharge machining (EDM),
electrochemical machining (ECM) or electrostream (ES) machining,
through tip cap 30 to provide communication to the interior of the
blade. The channels may assume other convenient configurations,
such as substantially rectantangular or substantially triangular.
The final configuration is not critical, as any arcuate
configuration of an appropriate area is satisfactory. Channels 6
may continue adjacent to the squealer tip wall 4 substantially
perpendicular to tip cap 30. Alternatively, channels 6 may be cast
into tip cap 30 and grooves may be cast along squealer tip wall 4.
Because of the small size of the channels 6 and the physical
effects of the drilling on the material, it is difficult to
maintain the true or preselected cross-sectional configuration, due
to localized recast metal. For example, when the preferred circular
cross section is fabricated into tip cap 30 and along adjacent
squealer tip wall 4, it is not uncommon for the cross section
locally to be oval or elliptical. A first channel end 8 terminates
at an exit orifice 10 at a surface 12 of tip cap 30. A second end
(not shown) of channel 6 provides a fluid communication to the
cooling circuits (not shown) contained internally within the
turbine engine blade, generally located below tip cap 30 in FIG.
2.
A bond coat 14 is then formed on tip cap 30 and squealer tip 2,
include squealer tip walls 4 and optionally on squealer tip
pressure side wall 31 and optionally on suction side (backside)
wall 33. The bond coat may be a diffusion aluminide bond coat, such
as a NiAl or PtAl bond coat or combinations thereof, or it may be
an additive bond coat of NiAl or Pt Al applied by well established
techniques, for example, CVD, VPA and PACH. Alternatively, the bond
coat 14 may be a MCrAl(X) additive layer where M is an element
selected from the group consisting of Fe, Co and Ni and
combinations thereof and (X) is an element selected from the group
of gamma prime formers, solid solution strengtheners, consisting
of, for example, Ta, Re and reactive elements, such as Y, Zr, Hf,
Si, and grain boundary strengtheners consisting of B, C and
combinations thereof. The MCrAl (X) is applied in the traditional
manner using well-known techniques, for example, physical vapor
deposition (PVD) processes such as electron beam (EB), ion-plasma
deposition, or sputtering, and deposition temperatures can be
1600.degree. F. or higher. Thermal spray processes such as air
plasma spray (APS), low pressure plasma spray (LPPS) or high
velocity oxyfuel (HVOF) spray can also be used. The channels
through tip cap 30 may be masked as is well known in the art to
prevent the holes from being filled by the applied bond coat metal,
if there is such a concern.
Bond coat 14 is formed to a thickness of about 0.0005" to about
0.010", preferably about 0.002" in thickness. As discussed in the
above referenced co-pending applications, when the bond coat 14 is
applied after the channels 6 have been drilled, bond coat 14 may
partially close the exit orifices 10 of the channels 6. When bond
coat 14 is applied first, followed by drilling of the small
channels 6, the potential problem of hole blockage by bond coat 14
is eliminated. However, application of bond coat 14 after drilling
of channels 6 is preferred, to allow bond coat 14 to partially
penetrate along the walls channels 6, thereby increasing adherence
of the subsequently applied ceramic top coat. Application of bond
coat 14 after drilling of channels 6 also provides a protective
environmental coating over the exposed substrate forming the
internal walls of channel 6.
After channels 6 are generated and the bond coat 14 is applied, a
generally porous TBC top coat 16 comprised of a porous
yttria-stabilized zirconia (YSZ), typically including about 7-9% by
weight yttrium, is applied on top of the bond coat 14. The porous
YSZ structure can be achieved, for example, by applying the YSZ
using PVD or plasma spray processes at temperatures in the range of
1600.degree.-1800.degree. F., which are lower than traditional YSZ
application temperatures of 1825.degree.-2150.degree. F. Other
methods may be utilized independent of the reduced temperature
techniques or in combination with the reduced temperature
techniques to achieve the porous YSZ structure. Alternatively, the
porous ceramic thermal barrier layer may be a porous Al.sub.2
O.sub.3, or other suitable oxides, such as, for example, zirconia
modified by other refractory oxides such as oxides formed from
Group IV, V and VI elements, or oxides modified by Lanthanide (Rare
Earth) Series elements such as La, Nd, Gd, Yb or other elements in
the series.
The porous TBC layer 16 having a first preselected porosity is
applied to a thickness of at least about 0.003", preferably about
0.010". When about 0.010" of porous TBC layer 16 is applied, the
channels 6 can be completely filled with TBC 16, while reduced
thicknesses are applied when the object or purpose is to only
partially fill exit orifice 10 of channels 6 with porous TBC
material. A porous TBC layer 16 applied to a thickness greater than
about 0.010" typically will completely span exit orifice 10 of
channels 6.
Applied on top of the porous TBC layer 16 is a dense, relative to
the porous TBC layer 16, ceramic layer 18 having a second,
preselected porosity or density. The dense ceramic layer 18 may be,
for example, YSZ, a thin layer of Al.sub.2 O.sub.3, or any other
suitable oxide, for example, zirconia modified by other refractory
oxides, such as oxides formed from Group IV, V and VI elements, or
oxides modified by Lanthanide (Rare Earth) Series elements such as
La, Nd, Gd, Yb or other elements in the series. Such oxide layers
may be deposited by physical deposition processes such as EB-PVD.
The method for forming a layer of a suitable oxide over dense
ceramic layer 18 is not restricted to PVD techniques, and other
suitable processes also may be employed. The dense ceramic layer is
applied to a thickness of about 0.002" to about 0.020". Although it
is possible to provide a thicker layer of dense ceramic, thicker
layers are undesirable because of the added weight. Preferably ,
the dense TBC layer is applied to a thickness in the range of
0.002"-0.003".
When bond coat has previously been applied, a conventional TBC
layer 20 optionally may be applied to wall 31 along concave
(pressure side) 22 and/or wall 33 along convex (suction side) 24
airfoil surfaces using deposition techniques similar to those used
to apply the porous layer 16. This TBC may have a third preselected
porosity or density. The density of this layer can be varied as
desired using well known methods for varying deposition densities
such as by varying the deposition temperature. The beneficial
cooling effects of the present invention extend the life of the TBC
by creating increased adherence of the pressure and suction side
TBC to the airfoil component due to introduction of cooling of the
airfoil in these regions, which heretofore has not been
incorporated into such blade designs.
Because the porous TBC layer 16 is processed to have a
predetermined porosity, cooling fluid, for example, cooling air, is
able to flow through the channels 6 and spread inside the porous
TBC layer 16. The dense ceramic layer 18 is much denser than the
porous TBC, inhibiting through passage of cooling fluid. Therefore,
the porous TBC layer 16 located between the bond coat 14 and the
dense ceramic layer 18 effectively forms a cooling channel
directing the flow of cooling air from the channel exit orifice 10
to the squealer tip 2. During engine operation, cooling air flows
between the bond coat 14 and the dense ceramic layer 18, and exits
to the tip 2 or any locations inside the tip cavity, eventually
discharging into the gas stream. In this manner, the TBC system has
both cooling and insulation purposes. Because the composition
and/or microstructure of the dense ceramic TBC layer 18 is
different from the porous TBC layer 16, its structure may be
controlled as required for specific applications, for example, for
"hotspots" located on the engine component.
In operation, the cooling fluid passes into the cooling channels 6
from the cooling circuit (not shown). As it reaches the exit
orifice 10, which is partially filled by the bond coat 14 and
partially or completely filled by the porous TBC layer 16, the
cooling fluid is diverted into the tortuous porosity that forms
passageways for the cooling fluid. As the cooling fluid traverses
through the porous TBC layer 16 to the dense ceramic layer 18, it
encounters more resistance than it would in passing through an
unobstructed orifice, and thus is further diverted across a larger
volume of airfoil as the fluid seeks the path of least resistance
through the porous TBC. The dense outer ceramic layer 18 provides
substantial resistance to the cooling fluid, thereby preventing
through-passage of the cooling fluid external to the airfoil.
Instead, dense ceramic layer 18 assists in directing the cooling
fluid along squealer tip walls 4 where it can readily exit into the
gas stream adjacent squealer tip 2. As the cooling fluid traverses
through the passageways, it removes heat from the adjoining TBC
through which it passes. The cooling fluid, which is at an elevated
temperature, ultimately is expelled, typically into the gas stream.
In this manner, the bond coat 14 is kept at a reduced temperature
through convection cooling. Utilizing the convection cooling
techniques and insulation provided by the present invention in
portions of airfoils that have not previously included these
features, a gas turbine engine is able to be operated at
temperatures hotter than those presently employed, with a resulting
increase in engine efficiencies.
Although the present invention has been described in connection
with specific examples and embodiments, those skilled in the art
will recognize that the present invention is capable of other
variations and modifications within its scope. These examples and
embodiments are intended as typical of, rather than in any way
limiting on, the scope of the present invention as presented in the
appended claims.
* * * * *